In a large frame heavy-duty industrial gas turbine engine, a hot gas stream generated in a combustor is passed through a turbine to produce mechanical work. The turbine includes one or more rows or stages of stator vanes and rotor blades that react with the hot gas stream in a progressively decreasing temperature. The efficiency of the turbine—and therefore the engine—can be increased by passing a higher temperature gas stream into the turbine. However, the turbine inlet temperature is limited to the material properties of the turbine, especially the first stage vanes and blades, and an amount of cooling capability for these first stage airfoils.
The first stage rotor blade and stator vanes are exposed to the highest gas stream temperatures, with the temperature gradually decreasing as the gas stream passes through the turbine stages. The first and second stage airfoils (blades and vanes), as well as shrouds and other structures, must be cooled by passing cooling air through internal cooling passages and discharging the cooling air through film cooling holes to provide a blanket layer of cooling air to protect the cooled surface from the hot gas stream.
The turbine rotor blades rotate within a cooled surface formed by a shroud which forms a gap with the blade tips. The shroud is formed of many segments secured within a ring carrier. A hot gas flow leakage that passes through the gap not only decreases the turbine efficiency but also creates hot spots on the shroud that result in erosion or other thermal induced damage for a short part life.
Current manufacturing processes for micro-channel intake ends and exhaust ends require high precision machining to align the ends with the correct channel that each end serves. These manufacturing limitations prevent flexibility in the size, cross sectional area and flow area of the intake and exhaust ends for metering of cooling air flow. With current tolerance stacking, micro-channels cannot be metered.